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Full text of "Aerodynamic characteristics of the NACA 747A315 and 747A415 airfoils from tests in the NACA two-dimensional low-turbulence pressure tunnel"

CB No. U125 



NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 



WARTIME REPORT 



ORIGINALLY ISSUED 

September 19^4 as 
Confidential Bulletin L4I25 



AERODYHAMIC CHAEACTEEISTICS OF THE NACA 7l*7A315 AKD 

747AUI5 AIRFOILS FROM TESTS IN TEE 
NACA TW0-DB4ENSI0NAL LOW-TURBULENCE PRESSURE TUNNEL 
By Albert E. von Doenhoff and Louis S. Stivers, Jr. 



Langley Memorial Aeronautical Laboratory 
Langley Field, 7a. 



NACA 



WASHINGTON 

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of 
advance research results to an authorized group requiring them for the war effort. They were pre- 
viously held under a security status but are now unclassified. Some of these reports were not tech- 
nically edited. All have been reproduced without change in order to expedite general distribution. 



156 DOCUMENTS DcPAKTfvibNT 



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in 2011 witli funding from 

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http://www.archive.org/details/aerodynamictestsOOIang 



1l^ 



?^ 



NACA CB No. L4I25 

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 



CONFIDENTIAL BULLETIN 



AERODYNAMIC CHARACTERISTICS OF THE NACA 747A315 AND 

747A415 AIRFOILS FROM TESTS IN THE 
NACA TWO-Dir^SNSIONAL LOW-TURBULENCE PRESSURE TLTNNSL 
By Albert E. von Doenhoff and Louis S. Stivers, Jr. 

SUM'IARY 



Two low-drag airfoils, the NACA 747A315 and the 
NACA 747A415, designed to have reduced pitching moments 
about the quarter-chord point and moderately high values 
of the design lift coefficient have been tested in the 
NACA two-dim.ensional low-turbulence pressure tunnel. 
Section lift, drag, and pitchlng-moment coefficients are 
presented for Reynolds numbers of 3 x 10^, 6 x 10^, and 
9 X 10 , together with section lift and section drag data 



for a Reynolds number of 6 x 10 
with roughened leading edges. 



6 



for the same airfoils 



A comparison of the characteristics, at a Reynolds 



number of 9 x IC*^ 



of the NACA 747A315 and NACA 747A415 



airfoils with characteristics of the N'ACA 652-415 airfoil 
is given in the following table: 



MCA 


Minimum 


Range of 


Section pitching- 


Maximum 


Critical Mach 


airfoil 


section 


section 


moment coeffi- 


section 


number at 


section 


drag 


lift coef- 


cient about 


lift 


design sectiun 




coef- 


ficient 


quarter chord at 


coef- 


lift coeffi- 




ficient 


for low 
drag 


design section 
lift coefficient 


ficient 


cient 


747A315 


0.0038 


0.22 to 0.62 


-0.017 


1.43 


0.626 


747A415 


.0041 


..32 to .'/2 


-.036 


1.50 


.612 


6 5- -41 5 


.0042 


.08 to .58 

1 


-.071 


1.6 2 


.641 
1 



2 CONFIDENTIAL NACA CB No. I.4I25 

INTRODUCTION 

The type of mean line usually used to camber the 
C-sevies airfoi],s presented In reference 1 led to rela- 
tively high pltching-monient coefficients about the 
quarter-chord point for a given design lift coefficient. 
For example, the measured pitching-moment coefficients 
for airfoils having a mean line of type a = 1.0 
cainbered for a design lift coefficient of 0.4 were 
approximately -0.070. This moment coefficient is some- 
v;hat higher than is desirs.ble for many applications. 

The FACA 7-series airfoils were derived in an attempt 
to obtain moderately high values of the design lift coef- 
ficient and to retain the low-drag characteristics of the 
6-serie3 airfoils, but with reduced pitching moments. 
The KACA 7-serles airfoils differ fromi the 6-series air- 
foils in that the 7-serles airfoils have a slightly modi- 
fied thiclTneos disuributlon and are comibined with mean 
lines in such a manner that more e:itensive regions of 
laminar flow are possible ove.v the lower surface than 
over the upper surface. The chordv/ise load distribution 
is €0 chosen that the main portion of the lift is carried 
by the forward part of the airfoil and the pitching 
mom.ents about the quarter-chord point are thus reduced. 



?he present report gives data for two ai^^foils of 
:y[)c-, de.-ignabi.d the NACA 747ii315 and the 



Thf 
this tyj)^ 

NACA 747A415 airfoils, from tests in the NACA two- 
dimensional low-turbulence pressure tiinnel (TDT) . Lift, 
drag, and pitching m-^m.ents of these sections were meas- 
ured for a range cf i^o^'nolds numbers from 3 to 9 x IC^. 
The effect of roughness at the leading edge on the lift 
and drag characteristics of the sections was determined 
at a Reynolds number of 6 x 10°, 



SY?,©OLS 

X distance from airfoil leading edge measured 

along chord line 

Cq section angle of attack 
c airfoil chord length 

COI.T'IDI]NTIAL 



NACA CB No. L4I25 CONFIDENTIAL 

r,^ section lift coefficient 

cj . design lift coefficient for mean line 

c,j section drag coefficient 



'm, 



=/4 



section pitching-moment coefficient about 
quarter-chord point 



V local velocity over airfoil surface 

V free-stream velocity 

a mean-line designation described in reference 1 

R Reynolds number' 

Mi^ critical Mach number 

DERIVATION 0? AIRF'OILS 



In the derivation of the NACA 747A315 airfoil, an 
attempt vi^as made to have uniform, load from the leading 
.edge to 0.4c back of the leading edge, to have- this load 
decrease linearly to zero at 0.7c, and to have zero load 
from. 0.7c to the trailing edge. This object was attained 
by combining a mean line of type a = 0.4 for a design 
lift coefficient of 0.763 with a mean line of type 
a = 0.7 for a design lift coefficient of -0.463. Ordi- 
nates and load distributions for these mean lines may be 
derived from, data presented in reference 1. 

In order to maintain a favorable pressure gradient 
to 0.7c along the lov;er surface of the NACA 747A315 air- 
foil, the resulting mean line was combined with a modi- 
fication of the NACA 64,2-015 airfoil section. The 
64,2-015 airfoil section was modified to reduce the slope 
of the pressure gradient from 0.4c to 0.7c. Figure 1 
rhov/s the pressure distribution over the modified sym- 
metrical NACA 64,2-015 airfoil section, together with 
the pressure distribution of the NACA 747A315 airfoil at 
the design lift coefficient. Table I gives the ordinates 
of the modified 64,2-015 airfoil section. Table II gives 
the ordinates of the 747A315 airfoil section. 



COITFIDSNTIAL 



CONFIDENTIAL NACA CB IJo . L4I2J 



The NACA 747A415 airfoil was obtained by combining 
the modified 64,2-015 airfoil section with the following 
mean-line combination: a = 0,4, c^. = 0.763; a = 0,7, 

c^. = -0.463; and a = 1.0, cj. = 0.100. Figure 2 

shovifs the theoretical pressure distribution of the 
NACA 747A415 airfoil at the design lift coefficient, 
together with that of the basic symmetrical section (the 
modified NACA 64,2-015 airfoil) . Ordlnates for the 
NACA 747A415 airfoil section are given in table III. 



AIRFOIL DESIGNATION 



The significance of the nujnbering system for these 
airfoils is explained by the following exam.ple: In the 
designation NACA 747A415, the first number "7" indicates 
the new series number; the second number "4" indicates 
the extent over the upper surface, in tenths of the 
chord from the leading edge, of the region of favorable 
pressure gradient at the der. ;.gn lift coefficient; the 
third number "7" indicates the extent over the lower 
surface, in tenths of the chord from the leading edge, 
of the region of favorable pressure gradient at the 
design lift coefficient. The significance of the last 
group of three numbers is the samie as for the previous 
6-series airfoils (reference 1); that is, the first 
number following the letter gives the design lift coef- 
ficient in tenths, and the last two num.bers give the 
airfoil thickness in percent of the chord. The letter 
"A", v;hich follows the first three numbers, is a serial 
letter to distinguish different airfoils having parame- 
ters that would correspond to the same numerical desig- 
nation. For example, a second airfoil having the same 
extent of favorable pressure gradient over the upper and 
lower surfaces, the sam.e design lift coefficient, and 
the same m.axlm.um thickness as the original airfoil but 
having a different mean-line combination and thickness 
distribution would have the serial letter "Fj." 



TEST PRCC5DURZ 



The models of the NACA 747A315 and 747A415 airfoil 
sections had a chord of 24 inches and a span of 35.5 inches. 
The methods of constructing and testing these models were 

CONFIDENT I AL 



>:ACA CB No. L4I2£ CONFIDENTIAL 



the ss.vre as those described in reference 1. The normal 
corrections for v/lnd- tunnel wall interference were made 
to the data obtained in the TDT accordin?; to the fol- 
lowinp forr^ulas, in which the primed quantities refer 
to the values measured in the wind tunnel: 



^o = 1-015 a„' 



c^ = 0.990 c^' 



c^, = 0.975 cj' 



^m^/4 ~ C-'.990 ^m^^^' 



In addition to tests at Reynolds numbers of 3, G, and 
9 X 10 in the sm.ooth condition, tests were made at a 
Reynolds ntimber of 6 x 10^ with the leading edpjes of 
the airfoils roughened. The roughness consisted of 
carborun.dum grains with a maximum diameter of about 
C.OIC to 0.015 inch. These grains were thinly sprinkled 
over the leading-ed;3;e pf^rtlon of the wing section 
covering a region of I'j- Inches from the leading edge 

o 

on the upper and lower surfaces across the span of the 
model. A thin coat of shellac was used to hold the 
grains on the surface. 

Since the presentation of the data in reference 1, 
certain changes have been made in the miethod of com- 
puting lift coefficients from tunnel data. More acc^orate 
factors have been derived, which give the proportion of 
lift actually transferred to the floor and the ceiling 
of the tunnel in the finite length covered by the floor 
and ceiling orifices. The revised factors result in a 
decrease in the slope of the lift curve and a decrease 
in the values of the maximum lift coefficient of approxi- 
mately 4 percent. In addition to the change in these 
factors, a correction for increased blocking effect at 
angles of attack in the neighborhood of maximum lift has 
been applied to the data presented herein. For the 
present data, this additional blocking correction derived 
from pressure measurements along the floor and ceiling of 
the tunnel resulted in a further reduction of the maxlrum 
lift coefficient by between 1 and 2 percent. 

CONFIDENTIAL 



CONFIDENTIAL NACA CB No. 14 125 



IffiSULTS AND DISCUSSION 



Section llf 
;.r, the TDT test^ 



t. drag, 
are p 
747A415 airfoils Ir figu 
each airfoil, the extent 
as large as would be exp 
percent thick 6-series a 
drag: range I'cr "both airf 
hneher than the design 1 
combination of mean line 
differtrce is probably d 
in the rrean-line theory 
velocity over the upper 
be snail conpared with t 



and p itching-moment data obtained 
resented for ^ the NACA 747A315 and 
res 5 and 4, respectively. For 

of the low-drag range is nearly 
ected froni the tests of a 15- 
irfoil . The center of the lov/- 
oils, however, is about 0.1c j 
ift coefficient given from the 
s used for each airfoil. Thi.? 
ue to the approximations involved 
in v/hich the increments in 
and lower surfaces are assumed to 
he f ree-3trea;.i \elocity. 



A comparison of the characteristics of the ITACA 747A515 
and the NACA 747A415 airfoils with those of the NACA 6 5o -415 

airfoil at a Re-fnclds number of 9 x 10^ is given in the 
table that follows. The dats for the NACA 652-415 airfoil 

given in the table were obtained from tests In the TDT and 
have been reduced in the saine manner as for the V-series 
airfoils. The values of critical Mach number M^, were 
obtained from the theoretical low- speed pressure distri- 
butions by using the chart presented on page 20 of the 
supplement to reference 1. 



: AOA 
airfoil 
section 


Qmm 


Range of 

cj for 
low drag 


design cj 


•"max 


Kq at 
design 


7li.7A315 
7i^7Ai-15 
652-U5 


0.0038 

.orUi 

.OCi+2 


0.22 to 0.62 
.32 to .72 
.03 to .S3 


-O.CI7 
-.036 
-.071 


i.Ii3 

l.SO 
1.62 


. 626 

.612 

.614 
1 



Langley Memorial Aeronautical Laboratory 

National Advisory Com^nittee for Aeronautics 
Langley Field, Va» 



CONFIDENTIAL 



NACA CB No. L4I25 COITFIDE^JTIAL 

REPSRSNCE 



1. Jacobs, Eastman N., Abbott, Ira E., and Davidson, 
Milton: Preliminary Lov;-Drag-Airf oil and Flap 
Data from Te?;ts at Large Reynolds Numbers and Low 
Turbulence, and Supplefnent . " NACA ACR, l^Iarch 1942, 



;onfide-?-:tial 



NACA CE No. L4I25 



GCKPIDErlTIAL 



3 



TABLE I.- MODIFIED NACA 64,2-015 AIRFOIL ORDINATES 
jJ3tations and ordinates given in percent of airfoil ch-irc 



Station 


Ordinate 








.5 


1.199 


.75 


1.435 


1.25 


1.801 


2.5 


2.462 


5.0 


3.419 


7.5 


4 . 143 


10 


4.743 1 


15 


5.684 


20 


6.364 


25 


6.898 


30 


7.253 


35 


7.454 


40 


7.494 


45 


7.316 


50 


7.003 


55 


6.584 


60 


6.064 


65 


5.449 


70 


4.738 


75 


3.921 


80 


3.020 


85 


2,086 


90 


1.193 


95 


.443 


100 





I.E. rad: 


.us; 1.544 



NATIONAL ADVISORY COMMITTEE FCR AERONAUTICS 



CONFIDENTIAL 



I'TACA GP No. L4I25 



CONFIDENT I AL 



TABLE II.- NACA 747A315 AIRFOIL ORDTKATES 
Ij.-i.-.ations and ordlnates given in percent of airfoil chordj 



Upper ? 


urf ace 


Lower cur 


'face 


Station 


Ordinate 


Station 


Ordinate 





n 
w 








.229 


1.305 


.771 


-1.031 


.449 


1.599 


1.051 


-1.207 


.911 


2.065 


1.589 


-1.473 


2.109 




2.391 


-1.027 


4.564 


4.254 


5.435 


-2.518 


7.053 


5.286 


7 . 347 


-2.952 


9.558 


6.140 


10.442 


-3.504 


14 . 599 


7.497 


15.401 


-3.843 


19.668 


8.503 


20.332 


-4.247 


24.758 


9.242 


25.242 


-4.546 


29.867 


9.731 


30.133 


-4.773 


35.001 


9 . 982 


34.999 


-4.926 


40.200 


9.962 


39.800 


-5.020 


45.375 


9.572 


44.625 


-5.040 


50.4 47 


8.964 


49.553 


-5.014 


55.463 


8.206 


54.537 


-4.930 


60.435 


1 . i^^4 


59.565 


-4.772 


Q5.366 


6 .355 


64.634 


-4.509 


70.241 


5.351 


69.759 


-4.110 ■ 


75.130 


4.336 


74.370 


-3.502 


80.073 


3.295 


79.927 


-2.743 


85.058 


2.257 


84.962 


-1.915 


90.016 


1.239 


89.984 


-1.097 


95.004 


.481 


94.996 


-.405 


100 





100 







L.E. radii 


is: 1.644 




Slope 


of radius thi 


•^ough L.iu. : 


0.232 



^"ATICNAL ADVISORY COMMITTEE FOR AERONAUTICS 



CONFIDENTIAL 



NACA CB No. L4I25 



CO:iFIDEI\TTIAL 



10 



TABLE III.- NACA 747A415 AIRFOIL ORDINATLS 
!_3tatlons and ordinates given in percent of airfoil cliordl 



Upper s 


urf ace 


Lov/er suj 


rf ace 


Station 


Ordinate 


Station 


Ordinate 











•0 


.183 


1.318 


.817 


-.994 


.398 


1.622 


1 . 102 


-1.160 


.852 


2.106 


1.648 


-1.406 


2.041 


3.016 


2.959 


-1.822 


4. 487 


4.411 


5.513 


-2.349 


6.972 


5.488 


8.028 


-2.730 


9.476 


6.390 


10 . 524 


-3.038 


14.521 


7.827 


15.479 


-3.501 


19.598 


8.897 


20.402 


-3.845 


24.698 


9 . 687 


25.302 


-4.0S5 


29.813 


10.216 


30 . 132 


-4.286 


34.964 


10.497 


35.036 


-4.'i,ll 


40.175 


10.499 


39.824 


-4.485 


45.364 


10.121 


44.636 


-4.493 


50.447 


9.516 


49.553 


-4.462 


55.474 


8.753 


54.526 


-4.581 


60.454 


7.859 


59.546 


-4.235 


65.393 


6.873 


64.607 


-3 . 992 • 


70.273 


5.838 


69.'?'27 


-3.622 


75.164 


4.733 


74.836 


-5. 053 


80 . 107 


3.692 


79.893 


-2.344 


85.066 


2 . 592 


84.934 


-1.578 


90.037 


1.546 


89.963 


-.828 


95.015 


.639 


94.985 


-.247 


100 





100 







L.E. radiuE 


3: 1.544 




Slope 


f radius thrc 


)ugh L.L. : C 


.274 



NATIONAL ADVISORY COMJGTTES FOR AERONAUTICS 



CONFIDENTIAL 



NACA CB No. L4I25 



CONFIDENTIAL 



Fig. 1 



2.0 



1.6 



i.U 



1.2 






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k, 


























/ 


^ 


^ 










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NACA 7li7A515 


/ 


/ 
















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V- 


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Lower Surface 


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A 




Modified NACA 6U, 2-015 






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figure 1.- Theoretical pressure distribution for NACA 7U7A315 and modifiad 
NACA 6I4., 2-015 airfolla. 



NACA CB No. L4I25 



CONFIDENTIAL 



Fig. 2 



2.0 



1.8 



1.6 



1.1; 



(v) 



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2 



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CONFIDENTIAL 



Vlgure 2.- Tneoretlcal pressure distribution for NACA 7U7*4l5 »nd modified 
NACA 614., 2-015 airfoils. 



NACA CB No. L4I25 



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NACA CB No. L4I25 



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UNIVERSITY OF FLORIDA 



262 08105 008 9 



UNIVERSITY OF FLORIDA 
DOCUMENTS DERARTMENT 
120 MARSTON SCIENCE UBRARY 
P.O. BOX 117011 
GAINESVILLE, FL 32611-7011 USA